The present invention relates generally to gas turbine engines, and, more specifically, to turbine rotor blades therein.
In a gas turbine engine, air is pressurized in a compressor and mixed with fuel in a combustor for generating hot combustion gases. A high pressure turbine (HPT) follows the combustor and extracts energy from the combustion gases for powering the compressor. A low pressure turbine (LPT) follows the HPT and extracts additional energy from the combustion gases for powering an upstream fan in an aircraft turbofan engine application, or powers an external drive shaft for marine and industrial applications.
The turbines are arranged in stages including a stationary turbine nozzle having a row of vanes which direct the combustion gases into a corresponding row of turbine rotor blades. Each vane has an airfoil configuration extending radially in span between inner and outer bands which bound the combustion gases.
Each turbine blade includes an airfoil extending radially in span from a root at an integral platform which in turn extends from an integral dovetail for mounting the blade in a corresponding dovetail slot in the perimeter of a supporting rotor disk. The platform defines the inner boundary for combustion gases, and the radially outer tip of the airfoil is spaced closely adjacent to a surrounding turbine shroud that defines the outer boundary for the combustion gases.
The corresponding airfoils of the vanes and blades in each turbine stage have generally concave pressure sides and generally convex suction sides extending axially in chord between opposite leading and trailing edges for efficiently turning the combustion gases and extracting energy therefrom during operation. The differently shaped opposite sides of the airfoils therefore effect different velocity and pressure distributions thereover, and correspondingly experience different heat loads from the combustion gases in highly complex three dimensional (3D) distributions.
The first stage turbine nozzle and blades first receive the combustion gases from the combustor and therefore have the greatest heat loads of the various turbine stages. Accordingly, the vanes and blades are typically cast from state-of-the-art superalloy metals which have enhanced strength at elevated temperature for maximizing the useful life thereof during operation.
The vane and blade airfoils are hollow and include corresponding internal cooling circuits therein which receive a portion of the pressurized air bled from the compressor for cooling thereof during operation. The internal cooling circuits typically include multiple radial channels defined by corresponding radial partitions bridging the pressure and suction sides of the airfoil, and those sides typically include radial rows or columns of film cooling holes extending transversely therethrough.
The cooling holes have various configurations and are typically tailored for the specific location of the airfoil from root to tip and between leading and trailing edges and on the opposite pressure and suction sides of the airfoil. For example, the leading edge of the airfoil first receives the hot combustion gases and typically has several columns of showerhead and gill holes for providing convection cooling through the sidewalls and external film cooling from the discharged cooling air.
The pressure and suction sides typically also include additional columns of film cooling holes for re-energizing the external film of cooling air as it flows downstream toward the trailing edge.
The typical film cooling hole is cylindrical and suitably drilled through the sidewall of the airfoil at a shallow inclination angle resulting in an oval inlet inside the airfoil and oval outlet on the external surface of the airfoil. Cooling air is discharged through the film cooling hole as a small jet that creates a thin film downstream therefrom for providing a thermally insulating layer of air outside the airfoil. The individual holes in the columns have a close spacing or pitch for maintaining lateral continuity of the cooling film.
The trailing edge of each airfoil may have a dedicated column of cooling holes located along the trailing edge itself between the pressure and suction sides, or commonly along the pressure side of the airfoil immediately upstream of the trailing edge for providing dedicated cooling of the thin trailing edge.
Since the turbine blade rotates during operation on the perimeter of the supporting rotor disk, it is subject to substantial centrifugal force which in turn creates centrifugal stress in the blade, and the combustion gases are subject to substantial radial forces as they flow generally downstream in the axial direction past the turbine airfoils. The rotating turbine airfoils therefore experience substantially different velocity and pressure distributions of the combustion gases as opposed to the stationary nozzle vanes.
In particular, the blade tips are bathed in the combustion gases not only along the pressure and suction sides, but also along the radially outer edge thereof as the combustion gases leak past the airfoil tips in the small clearance with the surrounding turbine shroud. The airfoil tips typically include small radial extensions of the pressure and suction sidewalls that define a squealer rib extending radially outwardly from the tip floor which encloses the internal cooling circuit of the blade.
The tip floor typically includes additional cooling holes that discharge cooling air into the small tip cavity bounded by the surrounding squealer rib. And, the pressure side of the airfoil may include another row of film cooling holes immediately below the squealer rib for local cooling of the airfoil tip itself.
Typical turbine vanes and blades have corresponding airfoil configurations which increase in width downstream from the leading edge to a maximum thickness closely adjacent thereto and then converge and taper in thickness to a thin trailing edge. The airfoils also extend in radial span between their inner and outer ends in the different configurations of the vanes and blades.
Correspondingly, the combustion gases discharged from the annular combustor have a center biased peak in temperature with corresponding pattern and profile factors varying in temperature circumferentially and radially. The combustion gases therefore introduce different heat loads three dimensionally over the turbine vanes and blades, which heat loads are additionally affected by the rotation of the turbine blades.
Accordingly, the prior art in turbine vane and blade cooling is replete with different configurations for vane and blade cooling. The internal cooling circuits of the airfoils have myriad configurations for distributing the limited cooling air for maximizing cooling effectiveness thereof over the different parts of the airfoil.
Blade cooling must be effected with minimum use of air bled from the compressor which bleed air correspondingly decreases performance and efficiency of the turbine engine. However, the airfoils must be adequately cooled for obtaining a long useful life before experiencing undesirable thermal distress.
The various cooling holes found for the turbine airfoils also have a myriad of configurations and patterns for cooling the different portions of the airfoil differently against the corresponding heat loads from the combustion gases. For example, the typical film cooling hole is a relatively simple, inclined cylindrical hole which may be economically formed by laser drilling.
However, the cylindrical hole emits a jet of spent cooling air that is subject to the varying differential pressure with the external combustion gases. Each film cooling hole must have a suitable backflow margin to prevent ingestion of the combustion gases into the airfoil, but the backflow margin cannot be excessive or the discharge jet will separate from the external surface and reduce efficiency of the film cooling.
Accordingly, a more complex film cooling hole has a diverging configuration for diffusing the discharge cooling air to reduce its velocity and promote improved film cooling. The typical diffusion film cooling hole typically requires electrical discharge machining (EDM) with a correspondingly shaped electrode that significantly increases the time and cost of manufacture.
Accordingly, diffusion film cooling holes are avoided when possible, and are typically used in isolated columns for enhanced film cooling. Diverging diffusion holes therefore join the other types of dedicated cooling holes in a turbine airfoil available to the blade designer during development.
Modern gas turbine engine turbine airfoil design is therefore quite mature and sophisticated and provides the designer with a multitude of options in developing a modern turbine airfoil for the hostile environment of the turbine stages in which combustion gas temperature is ever driven upward for increasing efficiency of the engine. The typical dilemma facing the designer is the selection of the fewest cooling holes with the simplest configuration for the different parts of the turbine airfoil for obtaining acceptable cooling thereof with minimal air bled from the compressor for maximizing airfoil life.
Modern gas turbine engines have benefited from the continual development of turbine airfoil cooling, with further improvements nevertheless continuing in small but significant changes. Modern turbine airfoils may achieve years of service with thousands of hours of operation in remarkably long service life before experiencing undesirable thermal distress and the need for replacement thereof.
For example, one modern turbofan engine has enjoyed decades of successful commercial public use in the United States and other countries, for powering commercial aircraft in flight. A first stage turbine rotor blade is found in this exemplary turbofan engine that has undergone continual development over the engine program, and itself has enjoyed many years of operation with thousands of hours of service without thermal distress.
However, this long useful life of actual turbine blades in extended service has led to the discovery of localized thermal distress at the end of that long life. This parent turbine blade has been in public use and on sale for many years in the U.S. and abroad, and forms the basis for further improvement thereof as described hereinbelow.
Accordingly, it is desired to provide a turbine airfoil having further improved cooling for addressing this recently discovered thermal distress and further increasing the useful life thereof.